def get_max_cl(Re, r):
"""
Analyze airfoil at a fixed Re,
changing aoa from 10 to 15 by 0.1
and returns cl, cd, aoa that makes maximum cl
"""
xf = XFoil()
if r <= 0.175:
xf.airfoil = naca6409
else:
xf.airfoil = naca2412
xf.Re = Re
xf.Re = Re
xf.max_iter = 200
xf.n_crit = 9.00
xf.xtr = [1.00, 1.00]
xf.M = 0
a_seq, cl_seq, cd_seq, cm_seq, cp_seq = xf.aseq(10, 15, 0.1)
# ignore nan by making it 0
cl_seq = np.nan_to_num(cl_seq)
# find the maximum cl
cl_maxi = np.max(cl_seq)
# index of the maximum cl
idx = np.argmax(cl_seq)
return round(cl_maxi, 2), round(a_seq[idx], 2), round(cd_seq[idx], 2)
def hello():
x = request.args.get('x')
y = request.args.get('y')
Re = float(request.args.get('Re'))
M = float(request.args.get('M'))
Alpha = float(request.args.get('Alpha'))
x = x.split()
y = y.split()
ctrlX = [float(ele) for ele in x]
ctrlY = [float(ele) for ele in y]
bezierX, bezierY = airfoil(ctrlX, ctrlY, 16)
xf = XFoil()
xf.Re = Re
xf.M = 0
xf.max_iter = 100
xf.airfoil = Airfoil(np.array(bezierX), np.array(bezierY))
aero = xf.a(Alpha)
xcp, cp = xf.get_cp_distribution()
y = savgol_filter(cp, 5, 2)
for i in range(30):
y = savgol_filter(y, 5, 2)
LD = aero[0] / aero[1]
vol = PolyArea(bezierX, bezierY)
print(len(xcp))
return jsonify(result=str(round(aero[0], 3)) + " " +
str(round(aero[1], 3)) + " " + str(round(aero[2], 3)) +
" " + str(round(LD, 2)) + " " + str(round(vol, 3)),
xcp=xcp.tolist(),
cp=y.tolist())
def get_cl(coord, xf=None, angle=5):
if xf is None:
xf = XFoil()
xf.print = False
xf.Re = 3e6
xf.max_iter = 100
datax, datay = coord.reshape(2, -1)
xf.airfoil = Airfoil(x=datax, y=datay)
c = xf.a(angle)
cl= c[0]
return cl
def evaluate(self, individual):
ratios = self.decoder(individual, self.code_division)
datlist_list = [fc.read_datfile(file) for file in self.datfiles]
datlist_shaped_list = [
fc.shape_dat(datlist) for datlist in datlist_list
]
newdat = fc.interpolate_dat(datlist_shaped_list, ratios)
foil_para = fc.get_foil_para(newdat)
datx = np.array([ax[0] for ax in newdat])
daty = np.array([ax[1] for ax in newdat])
newfoil = Airfoil(x=datx, y=daty)
mt, mta, mc, mca, s = foil_para
penalty = 0
for g, p in zip(self.gs, self.penalties):
if (not g):
penalty += p
xf = XFoil()
xf.airfoil = newfoil
xf.Re = self.re
xf.max_iter = 60
print(self.assigns, self.datfiles)
scope = locals()
exec(self.assigns, scope)
#----------------------------------
#目的値
#----------------------------------
try:
obj1, obj2, obj3 = [eval(o) for o in self.Os]
except IndexError as e:
obj1, obj2, obj3 = [1.0] * self.NOBJ
traceback.print_exc()
except SyntaxError as e:
raise ValueError("invalid objection")
if (np.isnan(obj1) or obj1 > 1):
obj1 = 1
if (np.isnan(obj2) or obj2 > 1):
obj2 = 1
if (np.isnan(obj3) or obj3 > 1):
obj3 = 1
obj1 += penalty
obj2 += penalty
obj3 += penalty
return [obj1, obj2, obj3]
def getCoefficients(naca,
reynolds=1e6,
iterations=20,
angle=10,
angle_step=.5):
xf = XFoil()
xf.naca(f"{naca:04d}")
xf.Re = reynolds
xf.max_iter = 20
a, cl, cd, cm, cp = xf.aseq(0, angle, .5)
ratio = cl / cd
max_idx = np.nanargmax(ratio)
return (ratio[max_idx], a[max_idx]), a, cl, cd, max_idx
def cl_visc_analysis(airfoil, cl_i, cl_f, cl_step, re, mach, max_iter, id):
"""Viscous analysis over range of lift coefficients."""
# Initializes airfoil and assigns NACA
xf = XFoil()
xf.naca(airfoil)
xf.max_iter = max_iter
xf.Re = re
xf.M = mach
# Collects values
a, cl, cd, cm, cp = xf.cseq(cl_i, cl_f, cl_step)
x, cp_0 = xf.get_cp_distribution()
# Plots all the data
plot(a, cl, cd, cm, cp, x, cp_0, id)
def alpha_visc_analysis(airfoil, alpha_i, alpha_f, alpha_step, re, mach,
max_iter, id):
"""Viscous analysis over range of angle of attacks."""
# Initializes airfoil and assigns NACA
xf = XFoil()
xf.naca(airfoil)
xf.max_iter = max_iter
xf.Re = re
xf.M = mach
# Collects values
a, cl, cd, cm, cp = xf.aseq(alpha_i, alpha_f, alpha_step)
x, cp_0 = xf.get_cp_distribution()
# Plots all the data
plot(a, cl, cd, cm, cp, x, cp_0, id)
def get_torque(angular_velocity):
torque_sum_small = 0
torque_sum_large = 0
w = angular_velocity
for key, value in dfdict.items():
value["blade_velocity"] = value['r_position'] * w
value["relative_velocity"] = round(
math.sqrt(value["blade_velocity"]**2 + value["wind_velocity"]**2),
2)
value["arctan"] = math.degrees(
math.atan2(value["wind_velocity"], value["blade_velocity"]))
aoa = round(value["arctan"] - value["pitch_angle"], 2)
value["angle_of_attack"] = aoa
re_n = round(
value["relative_velocity"] * value["chord_length"] / 0.00001511, 3)
value["Reynolds_number"] = re_n
xf = XFoil()
if key < 13:
xf.airfoil = naca6409
else:
xf.airfoil = naca2412
xf.Re = re_n
xf.max_iter = 100
xf.n_crit = 9.00
xf.xtr = [1.00, 1.00]
xf.M = 0
value["Cl"], value["Cd"], value["Cm"], value["Cp"] = xf.a(aoa)
force_reference = 0.5 * density * value["relative_velocity"]**2
if math.isnan(value["Cl"]):
value["torque"] = 0
else:
lift = value["Cl"] * force_reference * 0.0125 * value[
'chord_length']
drag = value["Cd"] * force_reference * 0.0125 * value[
'chord_length']
value["torque"] = value["r_position"] * (
lift * math.sin(math.radians(value["pitch_angle"])) -
drag * math.cos(math.radians(value["pitch_angle"])))
if key < 13:
torque_sum_small += value["torque"]
else:
pass
if key > 0:
torque_sum_large += value["torque"]
else:
pass
df2 = pd.DataFrame.from_dict(dfdict, orient="index")
df_collection.append(df2)
torque_sum_avg = 0.5 * (torque_sum_small + torque_sum_large)
return torque_sum_avg
def total_dict(angular_velocity):
torque_sum = 0
w = angular_velocity
for key, value in dfdict.items():
value["blade_velocity"] = value['r_position'] * w
value["relative_velocity"] = round(
math.sqrt(value["blade_velocity"]**2 + value["wind_velocity"]**2),
2)
value["arctan"] = math.degrees(
math.atan2(value["wind_velocity"], value["blade_velocity"]))
aoa = round(value["arctan"] - value["pitch_angle"], 1)
value["angle_of_attack"] = aoa
re_n = round(value["relative_velocity"] * value["chord_length"] /
0.00001511)
value["Reynolds_number"] = re_n
xf = XFoil()
if key < 13:
xf.airfoil = naca6409
else:
xf.airfoil = naca2412
xf.Re = round(re_n / 100) * 100
xf.max_iter = 200
xf.n_crit = 9.00
xf.xtr = [1.00, 1.00]
xf.M = 0
c_l, c_d, c_m, c_p = xf.a(aoa)
force_reference = 0.5 * density * value["relative_velocity"]**2
if math.isnan(c_l):
pass
else:
value["Cl"] = c_l
value["Cd"] = c_d
value["Cm"] = c_m
value["Cp"] = c_p
lift = c_l * force_reference * 0.0125 * value['chord_length']
drag = c_d * force_reference * 0.0125 * value['chord_length']
value["lift"] = lift
value["drag"] = drag
# value["torque"] = value["r_position"] * lift * math.sin(math.radians(value["pitch_angle"]))
torque = value["r_position"] * (
lift * math.sin(math.radians(value["pitch_angle"])) -
drag * math.cos(math.radians(value["pitch_angle"])))
value["torque"] = torque
torque_sum += torque
xf.reset_bls()
# detailed_df = pd.DataFrame.from_dict(dfdict, orient="index")
# print(detailed_df)
print(torque_sum, angular_velocity)
return dfdict, torque_sum
def evaluate(individual):
global code_division
#----------------------------------
#遺伝子にも続いて新翼型を生成
#----------------------------------
#遺伝子をデコード
ratios = decoder(individual, code_division)
#遺伝子に基づき翼型を混合して、新しい翼型を作る
datlist_list = [fc.read_datfile(file) for file in datfiles]
datlist_shaped_list = [fc.shape_dat(datlist) for datlist in datlist_list]
newdat = fc.interpolate_dat(datlist_shaped_list,ratios)
#翼型の形に関する情報を取得する
#foilpara == [最大翼厚、最大翼厚位置、最大キャンバ、最大キャンバ位置、S字の強さ]
foil_para = fc.get_foil_para(newdat)
#新しい翼型をAerofoilオブジェクトに適用
datx = np.array([ax[0] for ax in newdat])
daty = np.array([ax[1] for ax in newdat])
newfoil = Airfoil(x = datx, y = daty)
mt, mta, mc, mca, s = foil_para
#----------------------------------
#翼の形に関する拘束条件
#----------------------------------
penalty = 0
print('===================')
if(mc<0):
print("out of the border")
print("reverse_cmaber")
penalty -= mc
if(mt<0.08):
print("out of the border")
print("too_thin")
penalty += 0.08-mt
if(mt>0.11):
print("out of the border")
print("too_fat")
penalty += mt-0.11
#if(foil_para[4]>0.03):
# print("out of the border")
# print("peacock")
# print('===================')
# return (1.0+(foil_para[4]-0.03),)*NOBJ
if(mta<0.23):
print("out of the border")
print("Atama Dekkachi!")
penalty += 0.23 - mta
if(mta>0.3):
print("out of the border")
print("Oshiri Dekkachi!")
penalty += mta - 0.3
#----------------------------------
#新翼型の解析
#----------------------------------
xf = XFoil()
xf.airfoil = newfoil
#レイノルズ数の設定
xf.Re = 1.5e5
#境界要素法計算時1ステップにおける計算回数
xf.max_iter = 60
#座標整形
xf.repanel(n_nodes = 300)
xf.print = False
#計算結果格納
a, cl, cd, cm, cp = xf.cseq(0.4, 1.1, 0.1)
lr = [l/d for l, d in zip(cl,cd)]
#----------------------------------
#目的値
#----------------------------------
try:
#揚抗比の逆数を最小化
obj1 = 1/lr[1]
#揚抗比のピークを滑らかに(安定性の最大化)
maxlr = max(lr)
maxlr_index = lr.index(maxlr)
obj2 = abs(maxlr - lr[maxlr_index+1])
#下面の反りを最小化(製作再現性の最大化)
obj3 = foil_para[4]
except Exception as e:
obj1,obj2,obj3=[1.0]*NOBJ
traceback.print_exc()
if (np.isnan(obj1) or obj1 > 1):
obj1 = 1
if (np.isnan(obj2) or obj2 > 1):
obj2 = 1
if (np.isnan(obj3) or obj3 > 1):
obj3 = 1
obj1 += penalty
obj2 += penalty
obj3 += penalty
print("individual",individual)
print("evaluate",obj1,obj2,obj3)
print("max_thickness",foil_para[0])
print("at",foil_para[1])
print("max_camber",foil_para[2])
print("at",foil_para[3])
print("S",foil_para[4])
print('===================')
return [obj1, obj2, obj3]
if
c = (a_max * 8 * pi * r * math.sin(phi) ** 2) / ((1 - a_max) * prop_blades * c_l * math.cos(phi))
re = (v_r * c) / (nu_mars)
# XFOIL #######
from xfoil import XFoil
xf = XFoil()
# Import an airfoil
from xfoil.test import XXXXX
xf.airfoil = XXXXX
# Setting up the analysis parameters
xf.Re = re
xf.max_iter = 100
xf.M = 0.7
# Obtaining the angle of attack, lift coefficient, drag coefficient and momentum coefficient of the airfoil
a, cl, cd, cm = xf.aseq(0, 30, 0.5)
def evaluate(self,individual):
DELTA = 1e10
#----------------------------------
#遺伝子に基づいて新翼型を生成
#----------------------------------
#遺伝子に基づきスプライン翼型を作成
x = individual[:int(len(individual)/2)]
x.insert(0,1.0)
x.insert(int(len(x)/2)+1,0.0)
x.append(1.0)
y = individual[int(len(individual)/2):]
if not (all([u - d > 0 for u, d in zip(y[:int(len(y)/2)], y[int(len(y)/2):])]) or all([u - d < 0 for u, d in zip(y[:int(len(y)/2)], y[int(len(y)/2):])])):
print("crossed")
return [DELTA*10]*self.NOBJ
y.insert(0,0.0)
y.insert(int(len(y)/2)+1,0.0)
y.append(0.0)
newdat = fc.spline_foil(x, y, 200)
shape_dat = fc.shape_dat([[a, b] for a, b in zip(newdat[0][::-1], newdat[1][::-1])])
#翼型の形に関する情報を取得する
foil_para = fc.get_foil_para(shape_dat)
mt, mta, mc, mca, s, crossed, bd, bt, bc, smooth, td = foil_para
# mt: 最大翼厚(百分率)
# mta: 最大翼厚位置(百分率)
# mc: 最大キャンバー(百分率)
# mca: 最大きゃんばー位置(百分率)
# s: 翼型の下面における、最大y座標-最小y座標
# crossed: 翼型が交差しているならTrue,それ以外ならFalse
# bd: 翼型の粗さ(大きいほど粗い)
# bt: 翼厚分布の粗さ(大きいほど粗い)
# bc: キャンバー分布の粗さ(大きいほど粗い)
# smooth: 無視
# td: 翼厚分布
if crossed:
print("crossed_a")
return [DELTA*10]*self.NOBJ
else:
print("hi_a")
#新しい翼型をAerofoilオブジェクトに適用
datx = np.array(newdat[0][::-1])
daty = np.array(newdat[1][::-1])
newfoil = Airfoil(x = datx, y = daty)
#翼型の形に関する拘束条件
penalty = 0
if not all([t >= 0.0035 for t in td[10:80]]):
penalty += 100 * (sum([abs(t - 0.0035)*10 for t in td[15:85] if t - 0.0035 < 0]))
if not all([t <= 0.015 for t in td[:15]]):
penalty += 100 * (sum([abs(t - 0.015)*10 for t in td[:15] if t > 0.015]))
if mta > 0.4:
penalty += 100 * (mta - 0.4)
if mc < 0.0:
penalty += 100 * (-mc)
if datx[0] > 1.002 or datx[0] < 0.998:
print("invalid foil")
return [DELTA*10]*self.NOBJ
#----------------------------------
#新翼型の解析
#----------------------------------
try:
xf = XFoil()
#レイノルズ数の設定
xf.airfoil = newfoil
xf.Re = self.re
xf.print = False
xf.max_iter = 40
#xf.polar = "polar" + id
#境界要素法計算時1ステップにおける計算回数
#xf.repanel(n_nodes = 180)
#計算結果格納
#result = xf.OneAlpha()
cl, cd, cm, cp = xf.a(5.0)
#----------------------------------
#目的値
#----------------------------------
if cl >= 0:
obj1 = 1/cl
else:
obj1 = self.delta
obj2 = cd
except Exception as e:
obj1,obj2=[DELTA]*self.NOBJ
traceback.print_exc()
if (np.isnan(obj1)):
obj1 = DELTA
if (np.isnan(obj2)):
obj2 = DELTA
return [obj1 + penalty, obj2 + penalty]
#!/usr/bin/python3
from xfoil import XFoil
import matplotlib.pyplot as plt
xf = XFoil()
xf.load("revclarky.dat")
xf.Re = 1e5
xf.M = 0
xf.max_iter = 100
a, cl, cd, cm, cp = xf.aseq(-2, 2, 0.5)
plt.plot(a, cl)
plt.title("alfa vs cl")
plt.show()
#plt.plot(xf.airfoil.x,xf.airfoil.y)
#plt.axis('equal')
#plt.show()
def analyze_airfoil(x,
y_u,
y_l,
cl,
rey,
mach=0,
xf=None,
pool=None,
show_output=False):
"""
Analyze an airfoil at a given lift coefficient for given Reynolds and Mach numbers using XFoil.
Parameters
----------
x : array_like
Airfoil x-coordinates
y_u, y_l : array_like
Airfoil upper and lower curve y-coordinates
cl : float
Target lift coefficient
rey, mach : float
Reynolds and Mach numbers
xf : XFoil, optional
An instance of the XFoil class to use to perform the analysis. Will be created if not given
pool : multiprocessing.ThreadPool, optional
An instance of the multiprocessing.Threadpool class used to run the xfoil_worker. Will be created if not given
show_output : bool, optional
If True, a debug string will be printed after analyses. False by default.
Returns
-------
cd, cm : float or np.nan
Drag and moment coefficients of the airfoil at specified conditions, or nan if XFoil did not run successfully
"""
# If the lower and upper curves swap, this is a bad, self-intersecting airfoil. Return 1e27 immediately.
if np.any(y_l > y_u):
return np.nan
else:
clean_xf = False
if xf is None:
xf = XFoil()
xf.print = show_output
clean_xf = True
clean_pool = False
if pool is None:
pool = ThreadPool(processes=1)
clean_pool = True
xf.airfoil = Airfoil(x=np.concatenate((x[-1:0:-1], x)),
y=np.concatenate((y_u[-1:0:-1], y_l)))
xf.Re = rey
xf.M = mach
xf.max_iter = 100
xf.n_crit = 0.1
cd, cm = pool.apply(xfoil_worker, args=(xf, cl))
if clean_xf:
del xf
if clean_pool:
del pool
return cd, cm, None if clean_xf else xf
for v in zip(*rankedAerofoils_names):
print(str(header).format(*v))
span = 52
rho = 0.4135
velocity = 230
g = 9.81
reynolds_div_chord = 7331382
mean_chord = 3
#https://www.grc.nasa.gov/WWW/K-12/airplane/reynolds.html
massMin = 70000
massMax = 90000
massSpacing = 10000
xf = XFoil()
#Reynolds number, unit chord (see XFOIL docs UNITS section)
xf.Re = reynolds_div_chord * mean_chord
xf.max_iter = 100
print("Reynolds Number: {0}\nMax iterations: {1}\n".format(xf.Re, xf.max_iter))
cl, massList = liftCoCalc()
allAerofoils_obj = initialiseAerofoils()
runAnalysis()
print("\n")
rankedAerofoils_obj, rankedAerofoils_names = sortResults()
print("\n")
displayResults()
SPAN_POINTS_PER_METRE = 10
chord_points = int((END_CHORD - START_CHORD) * CHORD_POINTS_PER_METRE)
chord_increment = float((END_CHORD - START_CHORD) / chord_points)
span_points = int((END_SPAN - START_SPAN) * SPAN_POINTS_PER_METRE)
span_increment = float((END_SPAN - START_SPAN) / span_points)
print("Max iterations: {0}\nMass: {1}Kg".format(xf.max_iter, mass))
for j in range(chord_points + 1):
chord = START_CHORD + (j * chord_increment)
chord = round(chord, 2)
mac = START_CHORD + (j * chord_increment) * (8 / (3 * pi))
mac = round(mac, 2)
for i in range(span_points + 1):
span = START_SPAN + (i * span_increment)
span = round(span, 2)
#Reynolds number, unit chord (see XFOIL docs UNITS section)
xf.Re = reynolds_div_chord * mac
print("\nSpan: {0}m\tRoot Chord: {1}m\tReynolds: {2}".format(
span, chord, xf.Re))
cl = liftCoCalc()
Aerofoils_obj = initialiseAerofoils()
runAnalysis()
displayResults()
saveResults()
def evaluate(self,individual):
#----------------------------------
#遺伝子に基づいて新翼型を生成
#----------------------------------
#遺伝子をデコード
ratios = self.decoder(individual, self.code_division)
#遺伝子に基づき翼型を混合して、新しい翼型を作る
datlist_list = [fc.read_datfile(file) for file in self.datfiles]
datlist_shaped_list = [fc.shape_dat(datlist) for datlist in datlist_list]
newdat = fc.interpolate_dat(datlist_shaped_list,ratios)
#翼型の形に関する情報を取得する
#foilpara == [最大翼厚、最大翼厚位置、最大キャンバ、最大キャンバ位置、S字の強さ]
foil_para = fc.get_foil_para(newdat)
#新しい翼型をAerofoilオブジェクトに適用
datx = np.array([ax[0] for ax in newdat])
daty = np.array([ax[1] for ax in newdat])
newfoil = Airfoil(x = datx, y = daty)
mt, mta, mc, mca, s = foil_para
#----------------------------------
#翼の形に関する拘束条件
#----------------------------------
"""
penalty = 0
for g, p in zip(self.gs, self.penalties):
if(not g):
penalty += p
"""
penalty = 0
print('===================')
if(mc<0):
print("out of the border")
print("reverse_cmaber")
penalty -= mc
if(mt<0.08):
print("out of the border")
print("too_thin")
penalty += 0.08-mt
if(mt>0.11):
print("out of the border")
print("too_fat")
penalty += mt-0.11
#if(foil_para[4]>0.03):
# print("out of the border")
# print("peacock")
# print('===================')
# return (1.0+(foil_para[4]-0.03),)*NOBJ
if(mta<0.23):
print("out of the border")
print("Atama Dekkachi!")
penalty += 0.23 - mta
if(mta>0.3):
print("out of the border")
print("Oshiri Dekkachi!")
penalty += mta - 0.3
#----------------------------------
#新翼型の解析
#----------------------------------
xf = XFoil()
xf.airfoil = newfoil
#レイノルズ数の設定
xf.Re = self.re
#境界要素法計算時1ステップにおける計算回数
xf.max_iter = 60
#print("hi")
#print(vars)
#scope = locals()
#var0, var1, var2, var3, var4, var5, var6, var7 = [0 if var == None or var == '' else eval(var,scope) for var in self.vars]
#計算結果格納
a, cl, cd, cm, cp = xf.cseq(0.4, 1.1, 0.1)
lr = [l/d for l, d in zip(cl,cd)]
#----------------------------------
#目的値
#----------------------------------
"""
try:
obj1,obj2,obj3 = [eval(o) for o in Os]
except Exception as e:
obj1,obj2,obj3=[1.0]*self.NOBJ
traceback.print_exc()
"""
try:
#揚抗比の逆数を最小化
obj1 = 1/lr[1]
#揚抗比のピークを滑らかに(安定性の最大化)
maxlr = max(lr)
maxlr_index = lr.index(maxlr)
obj2 = abs(maxlr - lr[maxlr_index+1])
#下面の反りを最小化(製作再現性の最大化)
obj3 = s
except Exception as e:
obj1,obj2,obj3=[1.0]*self.NOBJ
traceback.print_exc()
if (np.isnan(obj1) or obj1 > 1):
obj1 = 1
if (np.isnan(obj2) or obj2 > 1):
obj2 = 1
if (np.isnan(obj3) or obj3 > 1):
obj3 = 1
obj1 += penalty
obj2 += penalty
obj3 += penalty
return [obj1, obj2, obj3]
fig, ax = plt.subplots(1, 1, figsize=(9, 8))
klist = [1]
for k in range(1):
print(k)
x = aifoils[k]['x'].values
y = aifoils[k]['y'].values
ang_low = -32
ang_high = 32
ang_spacing = 2
xfc = XFoil()
xfc.airfoil = Airfoil(x, y)
xfc.Re = re[k]
xfc.max_iter = 40
ac, clc, cdc, cmc, cpminc = xfc.aseq(ang_low, ang_high, ang_spacing)
df = aifoils_cfd[k].where(aifoils_cfd[k]['alpha [deg]'] >= aoa_l)
df = df.where(aifoils_cfd[k] <= aoa_h)
df = df.dropna() #.values
cl_panel = []
cdp_panel = []
for aoa in df['alpha [deg]']:
data_xy = aifoils[k].values
CL, CDP, Cp, pp = panel(data_xy[:, :], alfader=aoa)
cl_panel.append(CL)
cdp_panel.append(CDP)
def evaluate(self, individual):
#解析が発散した際の評価値
DELTA = 1e10
#----------------------------------
#遺伝子に基づいて新翼型を生成
#----------------------------------
#遺伝子をデコード
ratios = self.decoder(individual, self.code_division)
#遺伝子に基づき翼型を混合して、新しい翼型を作る
datlist_list = [fc.read_datfile(file) for file in self.datfiles]
datlist_shaped_list = [
fc.shape_dat(datlist) for datlist in datlist_list
]
newdat = fc.interpolate_dat(datlist_shaped_list, ratios)
#翼型の形に関する情報を取得する
mt, mta, mc, mca, s, crossed, bd, bt, bc, smooth, td = fc.get_foil_para(
newdat)
#新しい翼型をAerofoilオブジェクトに適用
datx = np.array([ax[0] for ax in newdat])
daty = np.array([ax[1] for ax in newdat])
newfoil = Airfoil(x=datx, y=daty)
#----------------------------------
#翼の形に関する拘束条件
#----------------------------------
penalty = 0
#キャンバに関する拘束条件
if (mc < 0):
penalty -= mc
#最大翼厚に関する拘束条件
if (mt < 0.08):
penalty += 0.08 - mt
if (mt > 0.11):
penalty += mt - 0.11
#最大翼厚位置に関する拘束条件
if (mta < 0.23):
penalty += 0.23 - mta
if (mta > 0.3):
penalty += mta - 0.3
#----------------------------------
#新翼型の解析
#----------------------------------
xf = XFoil()
xf.airfoil = newfoil
#レイノルズ数の設定
xf.Re = self.re
#境界要素法計算時1ステップにおける計算回数
xf.max_iter = 60
xf.print = False
#計算結果格納
a, cl, cd, cm, cp = xf.cseq(0.4, 1.1, 0.1)
#----------------------------------
#目的値
#----------------------------------
try:
#揚抗比の逆数を最小化
obj1 = 1 / lr[1]
#揚抗比のピークを滑らかに(安定性の最大化)
maxlr = max(lr)
maxlr_index = lr.index(maxlr)
obj2 = abs(maxlr - lr[maxlr_index + 1])
#下面の反りを最小化(製作再現性の最大化)
obj3 = s
except Exception as e:
obj1, obj2, obj3 = [DELTA] * self.NOBJ
traceback.print_exc()
if (np.isnan(obj1) or obj1 > 1):
obj1 = DELTA
if (np.isnan(obj2) or obj2 > 1):
obj2 = DELTA
if (np.isnan(obj3) or obj3 > 1):
obj3 = DELTA
obj1 += penalty
obj2 += penalty
obj3 += penalty
return [obj1, obj2, obj3]
def feature_xfoil(cst_u,
cst_l,
t,
Minf: float,
Re,
AoA,
n_crit=0.1,
fname='feature-xfoil.txt'):
'''
Evaluate by xfoil and extract features.
Inputs:
---
cst-u, cst-l: list of upper/lower CST coefficients of the airfoil. \n
t: airfoil thickness or None \n
Minf: free stream Mach number for wall Mach number calculation \n
Re, AoA (deg): flight condition (s), float or list, for Xfoil \n
n_crit: critical amplification ratio for transition in xfoil \n
fname: output file name. If None, then no output \n
### Dependencies: cst-modeling3d, xfoil
'''
from cst_modeling.foil import cst_foil
from xfoil import XFoil
from xfoil.model import Airfoil
#TODO: Build foil
#! 201 is the maximum amount of points that xfoil can handle
#! tail = 0.001 is to avoid point overlap
xx, yu, yl, t0, R0 = cst_foil(201, cst_u, cst_l, x=None, t=t, tail=0.001)
#! xfoil do not support leading edge of (0,0) on both upper and lower surface
x = np.array(list(reversed(xx[1:])) + xx[1:])
y = np.array(list(reversed(yu[1:])) + yl[1:])
foil = Airfoil(x, y)
#TODO: Xfoil
xf = XFoil()
xf.print = False
xf.airfoil = foil
xf.max_iter = 40
#* Transition by power law
xf.n_crit = n_crit
#TODO: Xfoil calculation
if not isinstance(Re, list):
Re = [Re]
AoA = [AoA]
n = len(Re)
for i in range(n):
xf.reset_bls()
if Re[i] is not None:
xf.Re = Re[i]
cl, cd, cm, cp = xf.a(AoA[i])
x, cp = xf.get_cp_distribution()
print(xf.Re, AoA[i], cl)
#* Extract features
fF = PhysicalXfoil(Minf, AoA[i], Re[i])
fF.setdata(x, y, cp)
fF.extract_features()
#* Output
if fname is None:
continue
if i == 0:
f = open(fname, 'w')
else:
f = open(fname, 'a')
f.write('\n')
f.write('%10s %15.6f \n' % ('Minf', Minf))
f.write('%10s %15.6f \n' % ('AoA', AoA[i]))
f.write('%10s %15.6f \n' % ('Re', Re[i] / 1e6))
f.write('%10s %15.6f \n' % ('CL', cl))
f.write('%10s %15.6f \n' % ('Cd', cd))
f.write('%10s %15.6f \n' % ('Cm', cm))
f.close()
fF.output_features(fname=fname, append=True)
from xfoil import XFoil from xfoil.test import naca0012 xf = XFoil() xf.airfoil = naca0012 xf.Re = 1e6 xf.max_iter = 40 a, cl, cd, cm, co = xf.aseq(-20, 20, 0.5)
b = Xwing_aero1[:, 1]
archive3 = open('lt.txt', 'w')
for line in range(len(aa)):
aux = [aa[line], b[line]]
string = str(aux).replace("[",
" ").replace(",",
" ").replace("]",
" ").replace("'", " ")
archive3.writelines(string + '\n')
aux.clear
# Calculates the aerodynamic coeff using Xfoil
archive3.close()
xf.load('lt.txt')
xf.Re = 1e3
xf.max_iter = 100
ann, cl, cd, cm, co = xf.aseq(1, 11, 1)
an = np.linspace(1, 10, 10)
desired = np.zeros(31)
desired[0:10] = cl
desired[10:20] = cd
desired[20:30] = cm
Surf, Iben, Itor, Xcg, Ycg = Struct.ArfoilProperties(1.0, Xwing_struct1)
desired[30] = Xcg
#2.Initialize Aero and Structural Models
#2.a. Aero Model
Elements, Xc, Xn, Xt, DL, A = Aero.InitializeAeroModel(Xwing_aero)
AoA, Vnorm = 0.0, 6.0
Vinf = np.array([np.cos(AoA / 57.3), np.sin(AoA / 57.3)]) * Vnorm #wind speed
foil = Airfoil(x, y)
#TODO: Xfoil
xf = XFoil()
xf.print = False
xf.max_iter = 40
xf.airfoil = foil
xf.xtr = [0.0, 0.0]
Minf = 0.2
AoA = 8.0
Re = 1e7
fname = 'feature-xfoil.txt'
xf.M = Minf
xf.Re = Re
cl, cd, cm, cp = xf.a(AoA)
x, cp = xf.get_cp_distribution()
with open(fname, 'w') as f:
f.write('%10s %15.6f \n'%('Minf', Minf))
f.write('%10s %15.6f \n'%('AoA', AoA))
f.write('%10s %15.6f \n'%('Re', Re/1e6))
f.write('%10s %15.6f \n'%('CL', cl))
f.write('%10s %15.6f \n'%('Cd', cd))
f.write('%10s %15.6f \n'%('Cm', cm))
fF = PhysicalXfoil(Minf, AoA, Re)
fF.setdata(x,y,cp)
fF.extract_features()